A gas turbine engine commonly includes an intake section, a compressor section, a combustor section, a turbine section, and an exhaust section. Due to the positioning of the turbine section immediately downstream of the combustor section, the air turbines and other turbine section components (e.g., turbine nozzles) are exposed to highly elevated temperatures during engine operation. Gas turbine engines are therefore often further equipped with a turbine cooling system to prevent overheating of the turbine section components by continually supplying cooling air thereto. The cooling air is bled from the gas turbine engine's compressor section and conducted through a network of cooling circuits, which directs the airflow over and around the combustion chamber before reintroducing the airflow into the turbine section. The turbine cooling system is typically passive in nature and relies upon the air pressure within the compressor section to drive airflow through the system's cooling circuits. In certain cases, the turbine cooling system may include a “Tangential On-Board Injection” or “TOBI” device, which injects the cooling airflow immediately upstream of the high pressure turbine while imparting the airflow with a tangential or swirling-type motion. In so doing, the TOBI device allows the bleed air to flow more easily into cooling channels provided in the rotating turbine thereby reducing parasitic pumping losses and providing lower cooling air temperatures and pressures to the turbine. The TOBI device thus serves as a means for producing a desired turbine cooling pressure and temperature that is lower than the maximum compressor exit condition.
To optimize the effectiveness of the turbine cooling system, the temperature of the air extracted from the compressor section is ideally as low as practical. At the same time, the pressure of the bleed air is preferably sufficiently high to create an adequate flow rate through the system's cooling circuits. Gas turbine engine platforms employing axial compressors typically have a relatively large number of compressor stages. As a result, it is typically relatively easy to select a compressor stage from which the turbine cooling system can bleed air that has a relatively low temperature while also having a sufficiently high pressure to satisfy the flow rate requirements of the cooling system. However, in the case of a gas turbine engine including a centrifugal compressor or impeller, it can be more difficult to extract air from the compressor section at a location that satisfies these competing criteria. When bled from a location near the inlet of the impeller, the temperature of the air is relatively low; however, so too is the air's pressure. Conversely, when bled from a location near the outlet of the impeller, the highly compressed airflow has a greatly elevated temperature and is generally undesirable for cooling purposes. The pressure level also tends to be much higher than necessary for adequate blade and disk cooling flow control and cooling passage pressurization. In addition, bleeding cooling air from the outlet of the impeller effectively wastes the energy expended to compress the airflow and consequently reduces the overall efficiency of the gas turbine engine.
It is thus desirable to provide embodiments of a gas turbine engine cooling system, such as a turbine cooling system, that enables the extraction of bleed air from an impeller at a location at which the temperature of the bleed air is relatively low, while the pressure of the bleed air is sufficiently high to satisfy the flow rate requirements of the cooling system. Ideally, embodiments of such a gas turbine engine cooling system would significantly reduce cooling circuit requirements, eliminate the need for a TOBI device, and decrease the overall part count, weight, and complexity of the cooling system as compared to a conventional turbine cooling system. It would also be desirable for embodiments of such a gas turbine engine cooling system to provide an improved impeller thermal stress gradient response and to improve overall system reliability by minimizing the amount of bleed air-entrained debris ingested by the cooling system. Finally, it would be desirable to provide embodiments of a method for producing such a gas turbine engine cooling system. Other desirable features and characteristics of the present invention will become apparent from the subsequent Detailed Description and the appended Claims, taken in conjunction with the accompanying Drawings and the foregoing Background.